Method of manufacturing a multiproperty rotor disk

ABSTRACT

A lightweight high temperature multiproperty rotor disk and method of manufacture for use in a gas turbine engine. The rotor disk of a nickel super alloy and having a plurality of lightweight high temperature single crystal blade attachment lugs bonded thereto. The single crystal high temperature rotor blade attachment lugs being cast in a substantially rectangular shape and bonded to the circumference of the disk. A method of manufacturing the rotor includes positioning the plurality of single crystal attachment lugs on a positioning ring that has locating nests formed thereon. Subsequently, the rotor disk is placed within the bonding ring and a restraining ring is placed around the entire subassembly. Upon subjecting the assembly to high temperature the rotor disk, having a coefficient of thermal expansion greater than the coefficient of thermal expansion of the restraining member, forces the components into a mating relationship. Upon completion of the bonding cycle, the restraining ring is removed from the assembly and the bonding ring is machined away.

This application is a division of application Ser. No. 08/568,986, filedDec. 7, 1994, now U.S. Pat. No. 5,609,471.

BACKGROUND OF THE INVENTION

The present invention relates generally to the design and constructionof a lightweight high temperature rotor disk system for a gas turbineengine. More particularly, the present invention has one form wherein aplurality of high strength blade attachment lugs, which are cast of asingle crystal alloy material, are bonded to a conventional powder metalnickel alloy disk. The high performance blade attachment lugs of thepresent invention have higher strength properties at elevatedtemperatures relative to conventional nickel alloy rims with noeffective increase in weight. Although the invention was developed foruse in gas turbine engines, certain applications may be outside of thisfield.

A gas turbine engine is typical of the type of turbomachinery in whichthe invention described herein may be advantageously employed. It iswell known that modern designers of gas turbine engines have generallyutilized an axial flow compressor for compressing air to the properpressure required for supporting the combustion of fuel in a combustionchamber. The high temperature exhaust gas exiting the combustion chamberprovides the working fluid for the turbine, which powers the axial flowcompressor. A power turbine that is driven by the flow of hightemperature gas is utilized to turn a propeller, fan or other device.Further, the high temperature gas may be used directly as a thrust forproviding motive power, such as in a turbine jet engine.

It is well known that the performance of gas turbine engines increasewith the increase in the operating temperature of the flow of hightemperature gas from the combustion chamber. A factor limiting theallowable temperature of the gaseous working flow from the combustionchamber is the capability of the various engine components to notdegrade when exposed to the high temperature gas flow. Further,maintainability of the gas turbine engine necessitates that the variouscomponents that are subjected to the high temperature gas flow must bereadily serviceable in order to minimize the down time of the gasturbine engine and the cost associated with repairs thereto. Varioustechniques have been utilized by engine designers to increase theallowable temperature of the gaseous working fluid, and to enhance theserviceability of engine components.

Gas turbine engine designers have generally sought to increase specificthrust, and reduce the specific fuel consumption in a gas turbineengine. However, associated with these parameters has generally been anincrease in turbine inlet temperature and compressor dischargetemperature. While the elevated fluid temperatures have increased theperformance of the gas turbine engines it has resulted in higher bladeplatform, stalk, and attachment temperatures which have been generallylowered by using additional compressor discharge cooling air. Further,the prior designers of gas turbine engines have sought to reduce theattachment temperature by increasing the stalk length of the blade toisolate the rim from the hot gaseous fluid flow path.

Cooling of the various components of the gas turbine engine ispreferably accomplished with a minimum amount of cooling fluid, sincethe cooling fluid is working fluid which has been extracted from thecompressor and its loss from the gas flow rapidly reduces engineefficiency. Therefore, the use of additional cooling fluid to cool theengine components increases the specific fuel consumption. Further, thedesign efforts to increase the temperature of the compressor dischargefluid, which is used to cool the blade platform, stalk and attachmentcompounds the cooling and specific fuel consumption issues.

With reference to FIG. 1, there is illustrated a conventional gasturbine blade `a` that is carried by a rotor disk (not illustrated). Gasturbine blade `a` has as principle regions an airfoil `b`, an attachmentportion `c` and a stalk `d` which extends between the attachment portion`c` and the airfoil `a`. A blade platform `e` is disposed above thestalk `d` and is designed to shield the turbine wheel from the flow ofhigh temperature gas from the combustion chamber. In many designs thestalk `d` function to elevate the platform `e` from the wheel in orderto minimize heat transfer to the wheel, and the correspondingdegradation thereof. One limitation generally associated with the priordesign is that the platform `e` and the stalk `d` increase the weight,centrifugal pull and attachment stress on the rotor system.

Heretofore, there has been a need for a lightweight high temperaturerotor disk for use in a gas turbine engine. The present inventionsatisfies this need in a novel and unobvious manner.

SUMMARY OF THE INVENTION

One form of the present invention contemplates a combination,comprising: a rotor disk; a plurality of single crystal attachment lugspositioned along the circumference of the rotor disk; and, a restrainingmember positioned adjacent the radial outward surface of the pluralityof attachment lugs, the restraining member having a coefficient ofthermal expansion less than the coefficient of thermal expansion of therotor disk.

Another form of the present invention contemplates a method for bondingsingle crystal attachment lugs to a rotor disk. The method comprising:providing a restraining member; positioning the single crystalattachment lugs along the circumference of the rotor disk; placing therestraining member around the radial outward surface of the attachmentlugs, and heating the assembly of the prior steps to join the attachmentlugs and the rotor disk.

One object of the present invention is to provide an improved rotorsystem for a gas turbine engine.

Related objects and advantages of the present invention will be apparentfrom the following description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a prior art gas turbine blade.

FIG. 2 is a perspective view of an aircraft having a gas turbine engineconnected thereto.

FIG. 3 is an enlarged partially fragmented side elevational view of thegas turbine engine of FIG. 2.

FIG. 4 is an illustrative partial side elevation sectional view of arotor with a high temperature turbine blade attachment lug according toone form of the present invention.

FIG. 5 is an exploded assembly view of a rotor assembly and fabricationtooling according to one form of the present invention.

FIG. 6 is a partial enlarged perspective view of the bond ring and hightemperature turbine blade attachment lug comprising a portion of theFIG. 5 rotor assembly and fabrication tooling.

FIG. 7 is a perspective view of the FIG. 5 rotor assembly andfabrication tooling with the fabrication tooling removed.

DESCRIPTION OF THE PREFERRED EMBODIMENT

For the purposes of promoting an understanding of the principles of theinvention, reference will now be made to the embodiment illustrated inthe drawings and specific language will be used to describe the same. Itwill nevertheless be understood that no limitation of the scope of theinvention is thereby intended, such alterations and furthermodifications in the illustrated device, and such further applicationsof the principles of the invention as illustrated therein beingcontemplated as would normally occur to one skilled in the art to whichthe invention relates.

Referring to FIGS. 2 & 3, there is illustrated an aircraft 10 having anaircraft flight propulsion engine 11. It is understood that an aircraftis generic and includes helicopters, tactical fighters, trainers,missiles and other related apparatuses. In the preferred embodiment theflight propulsion engine 11 includes a compressor 12, a combustor 13 anda power turbine 14. It is important to realize that there are amultitude of ways in which the components can be linked together.Additional compressors and turbines could be added with inner coolersconnecting between the compressors and reheat combustor chambers couldbe added between the turbines. Further, the gas turbine engine isequally suited to be used for industrial application. Historically therehas been wide spread application of industrial gas turbine engines, suchas pumping sets for gas and oil transmission lines, electricitygeneration and navel propulsion.

With reference to FIG. 3, there is illustrated an enlarged partiallyfragmented view of gas turbine engine 11 having a rotor system therein.A plurality of turbine blades 16 are insertably mounted to a disk 17that is affixed to a tubular shaft within the gas turbine engine 11. Aplurality of turbine vanes 16a forms a portion of the nozzle within thegas turbine engine. In the preferred embodiment the gas turbine blades16 and vanes 16a are unitary cast alloy structures produced by aprecision casting operation utilizing various super alloy compositions.Various types of super alloy compositions and manufactures of suchcompositions are known to those skilled in the art. Most super alloycompositions of interest are complicated mixtures of nickel, chromium,aluminum and other select elements. One technique for producing a castunitary turbine blade 16 or vane 16a having equiaxed, directionallysolidified, and single crystal alloy structures is disclosed in U.S.Pat. No. 5,295,530 to O'Connor which is incorporated by referenceherein. In the preferred embodiment the gas turbine blade 16 and vanes16a are of a single crystal alloy structure. It is understood that a gasturbine blade and a gas turbine vane are often referred to as anairfoil.

As used herein, a single crystal article is one in which substantiallyall of the article has a single crystallographic orientation through theload bearing portions, without the presence of high angle grainboundaries. A small amount of low angle grain boundaries such as tilt,or twist boundaries are permitted within such a single crystal articlebut are preferably not present. However, such low angle boundaries areoften present after solidification and formation of the single crystalarticle, or after some deformation of the article during creep or otherlight deformation process. Other minor irregularities are also permittedwithin the scope of the term "single crystal". For example, small areasof high angle grain boundaries may be formed in various portions of thearticle, due to the inability of the single crystal to grow perfectlynear corners and the like. Such deviations from a perfect singlecrystal, which are found in normal commercial production operations arewithin the scope of the term single crystal as used herein.

With reference to FIG. 4, there is illustrated a sectional view of arotor system 9. It is understood that the rotor disk 17 is symmetricalabout an axial centerline `x` of the gas turbine engine 11. In thepreferred embodiment the rotor disk 17 has a plurality ofcircumferentially spaced turbine blades 16 attached thereto. The rotordisk 17 includes a thickened rim section 18, a relatively thin websection 19 and a thickened bore section 20. A rotor disk may have othergeometric configurations, such as a uniform axial thickness from thebore section to the rim section. The plurality of turbine blades 16extend radially outward from the rim portion 18 and are carried by thedisk 17.

In the preferred embodiment the rotor system 9 is designed andmanufactured for installation into a turbine 14 of a flight propulsionengine 11. While the present invention has been described for use in aturbine, it is realized by those skilled in the art that it hasapplications in a compressor. In the preferred embodiment the disk 17 isformed of a powder metal nickel super alloy. In the most preferredembodiment the disk 17 is formed of AF2-1DA-6 powder metal materialwhich contains the following elements: about 0.04% carbon, about 12%chromium, about 10% cobalt, about 2.75% molybdenum, about 6.5% tungsten,about 2.8% titanium, about 1.5% tantalum, about 4.6% aluminum, about0.015% boron, about 0.10% zirconium, and about 59.695% nickel. It iscontemplated that the disk 17 can be formed of other materials havingsimilar properties to the materials cited herein.

In one form of the present invention a flow of cooling fluid from thecompressor 12 is discharged through a mechanical structure within thegas turbine engine 11 and directed to cool the rotor system 9. Acommonly owned U.S. patent application Ser. No. 08/509,777 filed Aug. 1,1995 discloses among other things a related cooling structure. U.S.patent application Ser. No. 08/509,777 is incorporated herein byreference.

Referring to FIG. 5, there is illustrated an exploded view of the rotorassembly and fabrication tooling 25. The rotor assembly and fabricationtooling 25 includes rotor disk 17, a plurality of blade attachment lugs22, a bond positioning ring 26, and a restraining ring 27. In one formof the present invention the disposable bond positioning ring 26 isfabricated from a conventional steel alloy. In contrast the restrainingring 27 is designed and manufactured as a reusable tool capable ofwithstanding the high temperature environment associated with bondingthe plurality of attachment lugs 22 to the circumference of disk 17.Bonding of the plurality of attachment lugs 22 to disk 17 isaccomplished through techniques known to those skilled in the art andinclude diffusion bond brazing in which a hyper eutectic composition ofnickel, chromium, and boron (Ni--Cr--B) is used.

In the preferred embodiment each of the plurality of blade attachmentlugs 22 are formed as a cast solid crystal article. It is preferred thatthe net cast shape be substantially rectangular. Each of the rectangularshaped lugs 22 including a planar outer surface 28 that is positionablewithin a nest 29 of the bond positioning ring 26. Other geometric shapesfor attachment lugs are contemplated herein, and the use of these lugsdepends upon the design of the specific rotor system. As previouslydiscussed herein, U.S. Pat. No. 5,295,530 to O'Connor discloses a methodfor producing a single crystal article. In the most preferred embodimentthe single crystal alloy material has the following composition: about6.5% chromium, about 9.0% cobalt, about 0.6% molybdenum, about 6.0%tungsten, about 1.0% titanium, about 6.5% tantalum, about 5.6% aluminum,about 61.8% nickel and about 3.0% rhenium. Material of this compositionis currently available from Cannon-Muskegan Corporation of Muskagan,Michigan and is sold under the trademark of CMSX-4®. It is understoodthat the lugs 22 can be formed of other single crystal alloy materialshaving similar properties to the materials recited herein.

Referring to FIGS. 5 and 6, there is illustrated a substantiallycylindrical bond positioning ring 26 that is utilized to receive,position and hold the plurality of blade attachment lugs 22 during therotor system manufacturing process. In the preferred embodiment the bondpositioning ring 26 is formed of a conventional steel alloy that willlose its structural integrity when subjected to the high temperaturesnecessary to diffusion bond the attachment lugs 22 to the disk 17. Bondpositioning ring 26 has formed thereon a plurality of circumferentiallyspaced nests 29. In the preferred embodiment the plurality of nests 29are uniformly spaced and have been machined on the inner cylindricalsurface of the bond positioning ring 26. Nests 29 in the preferredembodiment are defined as slots having a planar surface that receive theplanar surface 28 of the attachment lugs 22. It is understood that nestshaving other geometric shapes can be utilized to retain the bladeattachment lugs 22 on the bond positioning ring 26.

Blade attachment lugs 22 are fixedly attached to the inner cylindricalsurface of the bond positioning ring 26 at the respective nest 29locations. In the preferred embodiment the plurality of blade attachmentlugs 22 are attached at each nest location 29 by brazing. Othertechniques of affixing the blade attachment lugs 22 to the bondpositioning ring 26 such as tack welding are contemplated herein.Attachment of the blade attachment lugs 22 to the ring 26 maintainstheir respective orientation during the manufacturing process. In oneform of the present invention the plurality of nests 29 are oriented atan angle Θ to the axial centerline, and in one embodiment the angle Θ isabout ten degrees. It is understood that other angles are contemplatedherein, and are perimeters best left to the design of the specific rotorsystem. Parameters such as the number of blade attachment lugs, thenumber of blades, and the size of the hub will vary based upon specificrequirements of the rotor system.

The restraining ring 27 is utilized during the manufacturing process tominimize and/or eliminate outward radial movement of the componentsduring the diffusion bonding process. In the preferred embodiment therestraining ring 27 is a cylindrical member. However, the restrainingring 27 can have an outer surface of any shape so long as the member hasa substantially cylindrical internal surface 32. Prefereably, therestraining ring 27 has an inside diameter sized such that at roomtemperature it will slip over the bond positioning ring 26. Therefore atroom temperature when the restraining ring 27 is placed around the bondpositioning ring 26 there is some radial clearance between thesubstantially cylindrical internal surface 32 of the bond restrainingring 27 and the substantially cylindrical outer surface 33 of the bondpositioning ring 26.

In the preferred embodiment the plurality of blade attachment lugs 22have an axial length (l) which is greater than the axial length (p) ofthe bond positioning ring 26. The difference in length is designed tominimize the time associated with positioning the blade attachment lugs22 on the bond positioning ring 26. It is preferred that the axiallength of the restraining ring 27 be less than or equal to the axiallength of the bond joint between the attachment lugs 22 and the disk 17.In another form of the present invention the axial length of therestraining ring 27 is substantially equal to the axial length of thebond positioning ring 26.

With reference to FIGS. 4-7, a process associated with fabricating arotor system 9 will now be presented. Each of the plurality of bladeattachment lugs 22 are positioned in the nests 29 of the bondpositioning ring 26. Positioning of the blade attachment lugs 22involves locating the outer planar surface 28 adjacent the inner planarsurface 50 machined in the inner cylindrical surface of the bondpositioning ring 26. It is understood that lugs 22 in the preferredembodiment have four planar surfaces in the axial direction ofsubstantially equal size. However, in another form of the presentinvention the lugs have surfaces that are of unequal axial length. Asingle crystal attachment lug 22 is positioned on the disk such that thegrain structure is oriented in a substantially radial direction. In thepreferred embodiment the planar surface 28 of the blade attachment lugs22 corresponds to the planar surface 50 machined on ring 26. Havinglocated the lugs on the bond positioning ring 26 it is now necessary tofixedly attach them thereto. In the preferred embodiment this is done bya brazing process.

Subsequent to affixing the plurality of blade attachments lugs 22 to thebond positioning ring 26, this assembly is machined. The machining stepinvolves removing material from the inner surface 35 of each of theplurality of blade attachment lugs 22. Inner surfaces 35 of theplurality of blade attachment lugs 22 are machined to correspond to theouter geometric shape and diameter of the disk 17. It is understood thatby attaching the plurality of blade attachment lugs 22 to the bondpositioning ring 26 all of the attachment lugs can be machined at thesame time. Conventional manufacturing equipment such as a lathe and/orgrinder are contemplated to machine the inner surfaces 35 of theplurality of blade attachment lugs 22. The machining process produces acylindrical surface that is positionable at room temperature by a slipfit around disk 17.

After the inner surface 35 of the blade attachment lugs has beenmachined with sufficient radial clearance, it is positioned around theouter circumferential surface of the disk 17. Restraining ring 27 isthen slip fit, at room temperature, over the bond positioning ring 26.The unit is now assembled and ready for the bonding of the plurality ofblade attachment lugs 22 to the outer surface 31 of disk 17. It iscontemplated that the relative clearances and temperatures at which theprior assembly is undertaken can be adjusted as necessary by someone ofordinary skill in the art. Bonding of the lugs 22 to the disk 17 isaccomplished through techniques known to those skilled in the art andinclude diffusion bond brazing in which a hyper eutectic composition ofnickel, chromium, and boron (Ni--Cr--B) is used.

Subjecting the rotor assembly and fabrication tooling 25 to the heatassociated with diffusion bonding allows the differential thermalexpansion of the mating components to provide the pressure to hold andbond the blade attachment lugs 22 adjacent the outer surface 31 of thedisk 17. In the preferred embodiment the disk 17 has a coefficient ofthermal expansion that is greater than the coefficient of thermalexpansion of the restraining ring 27. The contribution to thermalexpansion associated with the single crystal blade attachment lug 22 isvery small. Further, upon reaching the diffusion bonding temperature thestructural integrity of the bond positioning ring 26 is degraded andit's expansion is null. Therefore, the thermal expansion of the disk 17forces the outer cylindrical surface 31 into the inner cylindricalsurface 35 of the blade attachment lugs 22 to facilitate bonding of thecomponents together. The restraining ring 27 eliminates and/or minimizesthe outward radial movement of the components and thereby a compressiveforce is applied to the components being bonded.

Upon removing the rotor assembly and fabrication tooling 25 from theheating source it is necessary to remove rings 26 and 27 from the rotorsystem. The outer restraining ring 27 is designed to be slid off of theassembly at room temperature. After the restraining ring 27 has beenremoved it is then necessary to remove the bond positioning ring 26 fromthe plurality of blade attachment lugs 22. Cutting of the bondpositioning ring 26 from the lugs 22 is accomplished by the use ofconventional machining processes. Techniques such as a grindingoperation and/or a turning operation are applicable.

In one form of the present invention the rotor disk 17 with lugs 22bonded thereto is subjected to a heat treating process. Heat treatingtechniques that are contemplated include an oil quench, a gas fan cool,an air cool, and a furnace cool. The heat treatment techniques areutilized in select situations to restore or enhance localized and/orbulk regions where material inconsistencies resulted from the diffusionbonding process.

Referring to FIG. 7, there is illustrated a perspective view of therotor system prior to the blade attachment lugs 22 being sized and priorto forming the blade interlocking arrangement. In one form of thepresent invention the blade attachment lugs 22 and the blade 16 willeach include an attachment portion that will interlock to connect theblade between a pair of the attachment lugs 22. After the requisitemachining operation the attachment portion of blades 16 are insertablebetween a pair of the blade attachment lugs 22 (not illustrated) Afirtree is the terminology recognized by those skilled in the art todescribe the interengagement of the attachment portions of the pluralityof lugs and blades. A firtree is utilized to more efficiently transferthe blade load to the firtree attachment. The present inventioncontemplated other methods of attaching the blades to the pair ofattachment lugs 22.

While the invention has been illustrated and described in detail in thedrawings and foregoing description, the same is to be considered asillustrative and not restrictive in character, it being understood thatonly the preferred embodiment has been shown and described and that allchanges and modifications that come within the spirit of the inventionare desired to be protected.

What is claimed is:
 1. A method for bonding single crystal attachmentlugs to a rotor disk, comprising:providing a restraining member;positioning the single crystal attachment lugs along the circumferenceof the rotor disk; placing the restraining member around the radialoutward surface of the attachment lugs; and heating the assembly of theprior steps to join the attachment lugs and the rotor disk.